1. Field of the invention
The invention concerns determining and controlling the attitude of a spin-stabilized satellite before it is transferred from an elliptical transfer orbit to a circular geostationary orbit by means of an apogee motor firing (AMF).
2. Description of the prior art
A satellite to be placed in a geostationary orbit is first launched from the ground by a launch vehicle such as the ARIANE rocket, for example, adapted to inject the satellite into an elliptical transfer orbit the perigee of which is at an altitude in the order of 200 km and the apogee of which is substantially at the altitude of the intended geostationary orbit (in the order of 36,000 kilometers). Then, when the satellite is passing through the apogee of the transfer orbit, an apogee motor firing is commanded during which an apogee motor on the satellite applies to it an impulse thrust with an orientation and amplitude adapted to transfer it substantially to its geostationary orbit.
At present a distinction is drawn between two types of launch procedure according to whether, when in the transfer orbit, the satellite is three-axis stabilized or spin-stabilized (rotates on itself) about an axis which is in practise the thrust axis of its apogee motor.
In the former case the orientation (or attitude) of the satellite is maintained constant, which entails frequent attitude correction maneuvers.
In the latter case, to which the invention relates, the orientation in space of the rotation axis of the satellite (here reference is also made to the "attitude" of the spinning satellite) varies only slightly along the transfer orbit. In practise, when the satellite is deployed by the launch vehicle an attempt is made to approximate as closely as possible the orientation (or attitude) required for firing the apogee motor and only slight attitude corrections are then made while the satellite is in its transfer orbit for fine adjustment of the actual attitude to the required attitude.
This orientation of the thrust axis is substantially perpendicular to the major axis of the elliptical transfer orbit.
It will be readily understood that the satellite can only be injected correctly into its geostationary orbit if, prior to the apogee motor firing, the thrust axis has been accurately oriented as required, failing which subsequent corrections which use up fuel will be necessary to reach the required orbit as well as can be achieved, which will reduce the residual quantity of fuel and therefore the service life of the satellite.
The orientation in space of the thrust axis (and therefore the attitude of the satellite) is determined by means of a set of terrestrial and solar sensors disposed on the satellite in an appropriate arrangement and the orientation is controlled and adjusted by thrusters provided on the satellite.
The solar and terrestrial sensors provide measurements of the angular offset between the direction in which they "see" the sun or the earth, respectively, and reference directions relating to the sensors. The determination of the orientation in space of an axis such as the thrust axis then follows preliminary recognition of the satellite-earth-sun plane.
Solar sensors are adapted to provide successive measurements throughout the transfer orbit except during periods of eclipse; as these are of limited duration and generally correspond to areas in the vicinity of the perigee, this is not disadvantageous.
Because of their positioning on the satellite relative to the thrust axis and because of their narrow field of view (.+-.1.degree. approximately), the terrestrial sensors are adapted to provide measurements over only a limited part of the transfer orbit in the vicinity of the apogee (and also in the vicinity of the perigee), where the earth actually enters the field of view of the terrestrial sensors, given the orientation in space of the satellite rotation axis (perpendicular to the major axis of the orbit).
A major difficultly in adjusting the attitude of the spinning satellite on initiating the apogee motor firing results from the fact that the satellite is usually at this time very near an imaginary line joining the earth and the sun. The result of this is significant uncertainty in the determination of the satellite-earth-sun plane and therefore in the determination of the orientation of the thrust axis.
This virtual alignment of the satellite between the earth and the sun at the time of the apogee motor firing results from the fact that satellite operators demand, for most currently available launch vehicles, and especially those designed to launch two satellites at one time, very narrow launch windows usually around midnight universal time. For the ARIANE rocket, for example, this launch window has a duration of 45 minutes and, depending on the launch date, ends at a time between 23 h 35 and 0 h 20.
The attitude of a spinning satellite before the apogee motor firing can therefore at present be determined only in the immediate vicinity of the apogee (given the positioning of the terrestrial sensors), in a satellite-earth-sun configuration that does not favor accurate measurement (because of the launch windows).
This state of affairs has led to various compromises or palliative measures being proposed:
seeking a derogation in respect of the launch procedures to avoid the times of year in which the launch window constraints are the most severe (equinoxes); the result of this is considerably reduced flexibility of use of the launch vehicle; PA1 accepting poor accuracy of attitude determination at the time of the apogee motor firing, at the cost of reducing the service life of the satellite; PA1 providing the satellite with additional terrestrial sensors positioned so that the earth lies in their field of view elsewhere than at the apogee of the transfer orbit; as these additional terrestrial sensors are used only during the transfer phase they represent a penalty in terms of cost and weight that is of no utility during the satellite's useful service life. PA1 the right ascension .beta. of the satellite is modified (C) to confer on it an intermediate attitude such that the earth is in the field of view of the terrestrial sensors for a position of the satellite in the transfer orbit offset at least 10.degree. from the apogee of the transfer orbit, PA1 the declination .alpha. of the rotation axis of the satellite is measured and then adjusted (C') to the value required for the apogee motor firing, PA1 maintaining this declination constant, the right ascension is adjusted (D) on the basis of the solar sensor measurements so as to bring the satellite into its final attitude, and PA1 the apogee motor firing is commanded. PA1 the right ascension .beta. differential between the initial attitude and the intermediate attitude is between 0.degree.-1.degree. and 20.degree., preferably around 10.degree.; PA1 in the intermediate attitude, the position of the satellite when the earth is in the field of view of the terrestrial sensors is approximately 10.degree. to 30.degree. ahead of the apogee; PA1 the apogee motor firing being scheduled for the fourth passage of the satellite through the apogee of the transfer orbit, the satellite is brought into its intermediate attitude before it passes the second time through the apogee of the transfer orbit; PA1 the satellite is brought into its intermediate attitude at a position of the satellite on its transfer orbit between 170.degree. and 10.degree. ahead of the apogee of the transfer orbit; PA1 the declination .alpha. of the satellite is adjusted at a position between 10.degree. and 30.degree. ahead of the apogee of the transfer orbit; PA1 the apogee motor firing being scheduled for the fourth passage of the satellite through the apogee of its transfer orbit, the declination .alpha. is adjusted before the satellite passes the second time through the apogee of its transfer orbit; PA1 the satellite is brought into its final attitude (C) at a position of the satellite on the transfer orbit between 10.degree. and 30.degree. ahead of the apogee of the transfer orbit; PA1 the apogee motor firing being scheduled for the fourth passage of the satellite through the apogee of its transfer orbit, the satellite is brought into its final attitude (D) before it passes the third time through the apogee of its transfer orbit; PA1 the apogee motor firing being scheduled for the fourth passage of the satellite through the apogee of its transfer orbit, a final adjustment (E) of the right ascension .beta. is commanded between 30.degree. and 10.degree. ahead of this fourth passage through the apogee of the transfer orbit.
An object of the invention is to alleviate these disadvantages by enabling accurate determination of the attitude of a spin-stabilized satellite before the apogee motor firing without any significant penalty in terms of weight or cost or satellite service life.